Structural configurations and cooling circuits in turbine blades

ABSTRACT

A turbine blade comprising an airfoil defined by a concave shaped pressure side outer wall and a convex shaped suction side outer wall that connect along leading and trailing edges and, therebetween, form a radially extending chamber for receiving the flow of a coolant. The turbine blade may further include: a rib configuration that partitions the chamber into radially extending flow passages that include a first flow passage and a second flow passage; and a crossover passage that fluidly connects an inlet formed in the first flow passage to an outlet formed in the second flow passage. The crossover passage may include a canted configuration relative to the second flow passage.

BACKGROUND OF THE INVENTION

This invention relates to turbine airfoils, and more particularly tohollow turbine airfoils, such as rotor or stator blades, having internalchannels for passing fluids such as air to cool the airfoils.

Combustion or gas turbine engines (hereinafter “gas turbines”) include acompressor, a combustor, and a turbine. As is well known in the art, aircompressed in the compressor is mixed with fuel and ignited in thecombustor and then expanded through the turbine to produce power. Thecomponents within the turbine, particularly the circumferentiallyarrayed rotor and stator blades, are subjected to a hostile environmentcharacterized by the extremely high temperatures and pressures of thecombustion products that are expended therethrough. In order towithstand the repetitive thermal cycling as well as the extremetemperatures and mechanical stresses of this environment, the airfoilsmust have a robust structure and be actively cooled.

As will be appreciated, turbine rotor and stator blades often containinternal passageways or circuits that form a cooling system throughwhich a coolant, typically air bled from the compressor, is circulated.Such cooling circuits are typically formed by internal ribs that providethe required structural support for the airfoil, and include multipleflow paths designed to maintain the airfoil within an acceptabletemperature profile. The air passing through these cooling circuitsoften is vented through film cooling apertures formed on the leadingedge, trailing edge, suction side, and pressure side of the airfoil.

It will be appreciated that the efficiency of gas turbines increases asfiring temperatures rise. Because of this, there is a constant demandfor technological advances that enable turbine blades to withstand everhigher temperatures. These advances sometimes include new materials thatare capable of withstanding the higher temperatures, but just as oftenthey involve improving the internal configuration of the airfoil so toenhance the blades structure and cooling capabilities. However, becausethe use of coolant decreases the efficiency of the engine, newarrangements that rely too heavily on increased levels of coolant usagemerely trade one inefficiency for another. As a result, there continuesto be demand for new airfoil designs that offer internal airfoilconfigurations and coolant circulation that improves coolant efficiency.

A consideration that further complicates design of internally cooledairfoils is the temperature differential that develops during operationbetween the airfoils internal and external structure. That is, becausethey are exposed to the hot gas path, the external walls of the airfoiltypically reside at much higher temperatures during operation than manyof the internal ribs, which, for example, may have coolant flowingthrough passageways defined to each side of them. In fact, a commonairfoil configuration includes a “four-wall” arrangement in whichlengthy inner ribs run parallel to the pressure and suction side outerwalls. It is known that high cooling efficiency can be achieved by thenear-wall flow passages that are formed in the four-wall arrangement,however, the outer walls experience a significantly greater level ofthermal expansion than the inner walls. This imbalanced growth causesstress to develop at the points at which the inner ribs and outer wallsconnect, which may cause low cyclic fatigue that can shorten the life ofthe blade. As such, the development of airfoil structures that usecoolant more efficiently while also reducing stress caused by imbalancedthermal expansion between internal and external regions remains asignificant technological industry objection.

BRIEF DESCRIPTION OF THE INVENTION

The present application thus describes a turbine blade comprising anairfoil defined by a concave shaped pressure side outer wall and aconvex shaped suction side outer wall that connect along leading andtrailing edges and, therebetween, form a radially extending chamber forreceiving the flow of a coolant. The turbine blade may further include:a rib configuration that partitions the chamber into radially extendingflow passages that include a first flow passage and a second flowpassage; and a crossover passage that fluidly connects an inlet formedin the first flow passage to an outlet formed in the second flowpassage. The crossover passage may include a canted configurationrelative to the second flow passage. In certain embodiments, thecrossover passage may be linear and configured so to form an acute angleof at least 20 degrees with a surface surrounding the outlet. The secondflow passage may be defined by radially extending sides that include: afirst side in which the outlet of the crossover passage is formed; and asecond side at which the crossover passage is aimed. The second side mayinclude a turbulator.

These and other features of the present application will become apparentupon review of the following detailed description of the preferredembodiments when taken in conjunction with the drawings and the appendedclaims.

BRIEF DESCRIPTION OF THE DRAWINGS

These and other features of this invention will be more completelyunderstood and appreciated by careful study of the following moredetailed description of exemplary embodiments of the invention taken inconjunction with the accompanying drawings, in which:

FIG. 1 is a schematic representation of an exemplary turbine engine inwhich certain embodiments of the present application may be used;

FIG. 2 is a sectional view of the compressor section of the combustionturbine engine of FIG. 1;

FIG. 3 is a sectional view of the turbine section of the combustionturbine engine of FIG. 1;

FIG. 4 is a perspective view of a turbine rotor blade of the type inwhich embodiments of the present invention may be employed;

FIG. 5 is a cross-sectional view of a turbine rotor blade having aninner wall or rib configuration according to conventional design;

FIG. 6 is a cross-sectional view of a turbine rotor blade having aninner wall configuration according to an embodiment of the presentinvention;

FIG. 7 is a cross-sectional view of the suction side of a turbine rotorblade having an inner wall configuration and flow passage communicationaccording to certain aspects of the present invention;

FIG. 8 is a cross-sectional view of the pressure side of a turbine rotorblade having an inner wall configuration and flow passage communicationaccording to certain aspects of the present invention; and

FIG. 9 is a sectional view of an outer wall of an airfoil according toan alternative embodiment of the present invention.

DETAILED DESCRIPTION OF THE INVENTION

As an initial matter, in order to clearly describe the current inventionit will become necessary to select certain terminology when referring toand describing relevant machine components within a gas turbine. Whendoing this, if possible, common industry terminology will be used andemployed in a manner consistent with its accepted meaning. Unlessotherwise stated, such terminology should be given a broadinterpretation consistent with the context of the present applicationand the scope of the appended claims. Those of ordinary skill in the artwill appreciate that often a particular component may be referred tousing several different or overlapping terms. What may be describedherein as being a single part may include and be referenced in anothercontext as consisting of multiple components. Alternatively, what may bedescribed herein as including multiple components may be referred toelsewhere as a single part. Accordingly, in understanding the scope ofthe present invention, attention should not only be paid to theterminology and description provided herein, but also to the structure,configuration, function, and/or usage of the component.

In addition, several descriptive terms may be used regularly herein, andit should prove helpful to define these terms at the onset of thissection. These terms and their definitions, unless stated otherwise, areas follows. As used herein, “downstream” and “upstream” are terms thatindicate a direction relative to the flow of a fluid, such as theworking fluid through the turbine engine or, for example, the flow ofair through the combustor or coolant through one of the turbine'scomponent systems. The term “downstream” corresponds to the direction offlow of the fluid, and the term “upstream” refers to the directionopposite to the flow. The terms “forward” and “aft”, without any furtherspecificity, refer to directions, with “forward” referring to the frontor compressor end of the engine, and “aft” referring to the rearward orturbine end of the engine. It is often required to describe parts thatare at differing radial positions with regard to a center axis. The term“radial” refers to movement or position perpendicular to an axis. Incases such as this, if a first component resides closer to the axis thana second component, it will be stated herein that the first component is“radially inward” or “inboard” of the second component. If, on the otherhand, the first component resides further from the axis than the secondcomponent, it may be stated herein that the first component is “radiallyoutward” or “outboard” of the second component. The term “axial” refersto movement or position parallel to an axis. Finally, the term“circumferential” refers to movement or position around an axis. It willbe appreciated that such terms may be applied in relation to the centeraxis of the turbine.

By way of background, referring now to the figures, FIGS. 1 through 4illustrate an exemplary combustion turbine engine in which embodimentsof the present application may be used. It will be understood by thoseskilled in the art that the present invention is not limited to thisparticular type of usage. The present invention may be used incombustion turbine engines, such as those used in power generation,airplanes, as well as other engine types. The examples provided are notmeant to be limiting unless otherwise stated.

FIG. 1 is a schematic representation of a combustion turbine engine 10.In general, combustion turbine engines operate by extracting energy froma pressurized flow of hot gas produced by the combustion of a fuel in astream of compressed air. As illustrated in FIG. 1, combustion turbineengine 10 may be configured with an axial compressor 11 that ismechanically coupled by a common shaft or rotor to a downstream turbinesection or turbine 13, and a combustor 12 positioned between thecompressor 11 and the turbine 13.

FIG. 2 illustrates a view of an exemplary multi-staged axial compressor11 that may be used in the combustion turbine engine of FIG. 1. Asshown, the compressor 11 may include a plurality of stages. Each stagemay include a row of compressor rotor blades 14 followed by a row ofcompressor stator blades 15. Thus, a first stage may include a row ofcompressor rotor blades 14, which rotate about a central shaft, followedby a row of compressor stator blades 15, which remain stationary duringoperation.

FIG. 3 illustrates a partial view of an exemplary turbine section orturbine 13 that may be used in the combustion turbine engine of FIG. 1.The turbine 13 may include a plurality of stages. Three exemplary stagesare illustrated, but more or less stages may be present in the turbine13. A first stage includes a plurality of turbine buckets or turbinerotor blades 16, which rotate about the shaft during operation, and aplurality of nozzles or turbine stator blades 17, which remainstationary during operation. The turbine stator blades 17 generally arecircumferentially spaced one from the other and fixed about the axis ofrotation. The turbine rotor blades 16 may be mounted on a turbine wheel(not shown) for rotation about the shaft (not shown). A second stage ofthe turbine 13 also is illustrated. The second stage similarly includesa plurality of circumferentially spaced turbine stator blades 17followed by a plurality of circumferentially spaced turbine rotor blades16, which are also mounted on a turbine wheel for rotation. A thirdstage also is illustrated, and similarly includes a plurality of turbinestator blades 17 and rotor blades 16. It will be appreciated that theturbine stator blades 17 and turbine rotor blades 16 lie in the hot gaspath of the turbine 13. The direction of flow of the hot gases throughthe hot gas path is indicated by the arrow. As one of ordinary skill inthe art will appreciate, the turbine 13 may have more, or in some casesless, stages than those that are illustrated in FIG. 3. Each additionalstage may include a row of turbine stator blades 17 followed by a row ofturbine rotor blades 16.

In one example of operation, the rotation of compressor rotor blades 14within the axial compressor 11 may compress a flow of air. In thecombustor 12, energy may be released when the compressed air is mixedwith a fuel and ignited. The resulting flow of hot gases from thecombustor 12, which may be referred to as the working fluid, is thendirected over the turbine rotor blades 16, the flow of working fluidinducing the rotation of the turbine rotor blades 16 about the shaft.Thereby, the energy of the flow of working fluid is transformed into themechanical energy of the rotating blades and, because of the connectionbetween the rotor blades and the shaft, the rotating shaft. Themechanical energy of the shaft may then be used to drive the rotation ofthe compressor rotor blades 14, such that the necessary supply ofcompressed air is produced, and also, for example, a generator toproduce electricity.

FIG. 4 is a perspective view of a turbine rotor blade 16 of the type inwhich embodiments of the present invention may be employed. The turbinerotor blade 16 includes a root 21 by which the rotor blade 16 attachesto a rotor disc. The root 21 may include a dovetail configured formounting in a corresponding dovetail slot in the perimeter of the rotordisc. The root 21 may further include a shank that extends between thedovetail and a platform 24, which is disposed at the junction of theairfoil 25 and the root 21 and defines a portion of the inboard boundaryof the flow path through the turbine 13. It will be appreciated that theairfoil 25 is the active component of the rotor blade 16 that interceptsthe flow of working fluid and induces the rotor disc to rotate. Whilethe blade of this example is a turbine rotor blade 16, it will beappreciated that the present invention also may be applied to othertypes of blades within the turbine engine 10, including turbine statorblades 17. It will be seen that the airfoil 25 of the rotor blade 16includes a concave pressure side outer wall 26 and a circumferentiallyor laterally opposite convex suction side outer wall 27 extendingaxially between opposite leading and trailing edges 28, 29 respectively.The sidewalls 26 and 27 also extend in the radial direction from theplatform 24 to an outboard tip 31. (It will be appreciated that theapplication of the present invention may not be limited to turbine rotorblades, but may also be applicable to stator blades. The usage of rotorblades in the several embodiments described herein is exemplary unlessotherwise stated.)

FIG. 5 shows an internal wall construction as may be found in a rotorblade airfoil 25 having a conventional design. As indicated, the outersurface of the airfoil 25 may be defined by a relatively thin pressureside outer wall 26 and suction side outer wall 27, which may beconnected via a plurality of radially extending and intersecting ribs60. The ribs 60 are configured to provide structural support to theairfoil 25, while also defining a plurality of radially extending andsubstantially separated flow passages 40. Typically the ribs 60 extendradially so to partition the flow passages over much of the radialheight of the airfoil 25, but, as discussed more below, the flow passagemay be connected along the periphery of the airfoil so to define acooling circuit. That is, the flow passages 40 may fluidly communicateat the outboard or inboard edges of the airfoil 25, as well as via anumber of smaller crossover passages or impingement apertures (notshown) that may be positioned therebetween. In this manner certain ofthe flow passages 40 together may form a winding or serpentine coolingcircuit. Additionally, film cooling ports (not shown) may be includedthat provide outlets through which coolant is released from the flowpassages 40 onto the outer surface of the airfoil 25.

The ribs 60 may include two different types, which then, as providedherein, may be subdivided further. A first type, a camber line rib 62,is typically a lengthy rib that extends in parallel or approximatelyparallel to the camber line of the airfoil, which is a reference linestretching from the leading edge 28 to the trailing edge 29 thatconnects the midpoints between the pressure side outer wall 26 and thesuction side outer wall 27. As is often the case, the conventionalconfiguration of FIG. 5 includes two camber line ribs 62, a pressureside camber line rib 63, which also may be referred to as the pressureside inner wall given the manner in which it is offset from and close tothe pressure side outer wall 26, and a suction side camber line rib 64,which also may be referred to as the suction side inner wall given themanner in which it is offset from and close to the suction side outerwall 27. As mentioned, this type of design is often referred to ashaving a “four-wall” configuration due to the prevalent four main wallsthat include the two sidewalls 26, 27 and the two camber line ribs 63,64. It will be appreciated that the outer walls 26, 27 and the camberline ribs 62 are cast as integral components.

The second type of rib is referred to herein as a traverse rib 66.Traverse ribs 66 are the shorter ribs that are shown connecting thewalls and inner ribs of the four-wall configuration. As indicated, thefour walls may be connected by a number of the traverse ribs 66, whichmay be further classified according to which of the walls each connects.As used herein, the traverse ribs 66 that connect the pressure sideouter wall 26 to the pressure side camber line rib 63 are referred to aspressure side traverse ribs 67. The traverse ribs 66 that connect thesuction side outer wall 27 to the suction side camber line rib 64 arereferred to as suction side traverse ribs 68. Finally, the traverse ribs66 that connect the pressure side camber line rib 63 to the suction sidecamber line rib 64 are referred to as center traverse ribs 69.

In general, the purpose of four-wall internal configuration in anairfoil 25 is to provide efficient near-wall cooling, in which thecooling air flows in channels adjacent to the outer walls 26, 27 of theairfoil 25. It will be appreciated that near-wall cooling isadvantageous because the cooling air is in close proximity of the hotouter surfaces of the airfoil, and the resulting heat transfercoefficients are high due to the high flow velocity achieved byrestricting the flow through narrow channels. However, such designs areprone to experiencing low cycle fatigue due to differing levels ofthermal expansion experienced within the airfoil 25, which, ultimately,may shorten the life of the rotor blade. For example, in operation, thesuction side outer walls 27 thermally expands more than the suction sidecamber line rib 64. This differential expansion tends to increase thelength of the camber line of the airfoil 25, and, thereby, causes stressbetween each of these structures as well as those structures thatconnect them. In addition, the pressure side outer wall 26 alsothermally expands more than the cooler pressure side camber line rib 63.In this case, the differential tends to decrease the length of thecamber line of the airfoil 25, and, thereby, cause stress between eachof these structures as well as those structures that connect them. Theoppositional forces within the airfoil that, in the one case, tends todecrease the airfoil camber line and, in the other, increase it, canlead to further stress concentrations. The various ways in which theseforces manifest themselves given an airfoil's particular structuralconfiguration and the manner in which the forces are then balanced andcompensated for becomes a significant determiner of the part life of therotor blade 16.

More specifically, in a common scenario, the suction side outer wall 27tends to bow outward at the apex of its curvature as exposure to thehigh temperatures of the hot gas path cause it to thermally expand. Itwill be appreciated that the suction side camber line rib 64, being aninternal wall, does not experience the same level of thermal expansionand, therefore, does not have the same tendency to bow outward. Thecamber line rib 64 then resists the thermal growth of the outer wall 27.Because conventional designs have camber line ribs 62 formed with stiffgeometries that provide little or no compliance, this resistance and thestress concentrations that result from it can be substantial.Exacerbating the problem, the traverse ribs 66 used to connect thecamber line rib 62 to the outer wall 27 are formed with linear profilesand generally oriented at right angles in relation to the walls thatthey connect. This being the case, the traverse ribs 66 operate tobasically hold fast the “cold” spatial relationship between the outerwall 27 and the camber line rib 64 as the heated structures expand atsignificantly different rates. Accordingly, with little or no “give”built into the structure, conventional arrangements are ill-suited atdefusing the stress that concentrates in certain regions of thestructure. The differential thermal expansion bus results in low cyclefatigue issues that shorten component life.

Many different internal airfoil cooling systems and structuralconfigurations have been evaluated in the past, and attempts have beenmade to rectify this issue. One such approach proposes overcooling theouter walls 26, 27 so that the temperature differential and, thereby,the thermal growth differential are reduced. It will be appreciated,though, that the way in which this is typically accomplished is toincrease the amount of coolant circulated through the airfoil. Becausecoolant is typically air bled from the compressor, its increased usagehas a negative impact on the efficiency of the engine and, thus, is asolution that is preferably avoided. Other solutions have proposed theuse of improved fabrication methods and/or more intricate internalcooling configurations that use the same amount of coolant, but use itmore efficiently. While these solutions have proven somewhat effective,each brings additional cost to either the operation of the engine or themanufacture of the part, and does nothing to directly address the rootproblem, which is the geometrical deficiencies of conventional design inlight of how airfoils grow thermally during operation.

The present invention generally teaches certain curving or bubbled orsinusoidal or wavy internal ribs (hereinafter “wavy ribs”) thatalleviate imbalanced thermal stresses that often occur in the airfoil ofturbine blades. Within this general idea, the present applicationdescribes several ways in which this may be accomplished, which includewavy camber line ribs 62 and/or traverse ribs 66, as well as certaintypes of angled connections therebetween. It will be appreciated thatthese novel configurations—which, as delineated in the appended claims,may be employed separately or in combination—reduce the stiffness of theinternal structure of the airfoil 25 so to provide targeted flexibilityby which stress concentrations are dispersed and strain off-loaded toother structural regions that are better able to withstand it. This mayinclude, for example, off-loading to a region that spreads the strainover a larger area, or, perhaps, structure that offloads tensile stressfor a compressive load, which is typically more preferable. In thismanner, life-shortening stress concentrations and strain may be avoided.

FIG. 6 provides cross-sectional views of a turbine rotor blade 16 havingan inner wall configuration according to embodiments of the presentinvention. Specifically, as aspect of the present invention involves theconfiguration of ribs 60 that are typically used as both structuralsupport as well as partitions that divide hollow airfoils 25 intosubstantially separated radially extending flow passages 40 that may beinterconnects as desired to create cooling circuits. These flow passages40 and the circuits they form are used to direct a flow of coolantthrough the airfoil 25 in a particular manner so that its usage istargeted and more efficient. Though the examples provided herein areshown as they might be used in a turbine rotor blades 16, it will beappreciated that the same concepts also may be employed in turbinestator blades 17. In one embodiment, the rib configuration of thepresent invention includes a camber line rib 62 having a wavy profile.(As used herein, the term “profile” is intended to refer to the shapethe ribs have in the cross-sectional views of FIG. 6.) A camber line rib62, as described above, is one of the longer ribs that typically extendfrom a position near the leading edge 28 of the airfoil 25 toward thetrailing edge 29. These ribs are referred to as “camber line ribs”because the path they trace is approximately parallel to the camber lineof the airfoil 25, which is a reference line extending between theleading edge 28 and the trailing edge 29 of the airfoil 25 through acollection of points that are equidistant between the concave pressureside outer wall 26 and the convex suction side outer wall 27. Accordingto the present application, a “wavy profile” includes one that isnoticeably curved and sinusoidal in shape, as indicated. In other words,the “wavy profile” is one that presents a back-and-forth “S” profile, asindicated in FIG. 6.

The segment or length of the camber line rib 62 that is configured withthe wavy profile may vary depending on design criteria. In the providedexamples the wavy camber line rib 62 typically stretches from a positionnear the leading edge 28 of the airfoil 25 to a position that is beyondthe midpoint of the camber line of the airfoil 25. It will beappreciated that the wavy portion of the camber line rib 62 may beshorter in length while still providing the same types of performanceadvantages discussed herein. The number of curves as well as the lengthof the wavy segment of the camber line rib 62 may be varied to achievethe best results. In certain embodiments, the wavy camber line rib 62 ofthe present invention is defined by the number of completeback-and-forth “S” shapes it contains. In a preferred embodiment of thistype, the wavy camber line rib 62 includes at least one continuousback-and-forth “S” shape. In another embodiment, the wavy camber linerib 62 includes at least two consecutive and continuous back-and-forth“S” shapes. In regard to overall length, the wavy segment of the camberline rib 62 may extend for a substantial portion of the length of thecamber line of the airfoil 25. For example, as shown in FIG. 6, in apreferred embodiment, the wavy portion of the camber line rib 62 is over50% of the length of the camber line of the airfoil 25. In other words,the wavy portion of the camber line rib 62 originates near the leadingedge 28 of the airfoil 25 and extend rearward and well beyond the apexof the curvature of the airfoil 25. It will be appreciated that shorterlengths also may be employed with performance benefits, such as wavyportions of at least 25% length of the camber line rib 62.

It will be appreciated that, given its winding profile, a wavy camberline rib 62 traces a path that varies in its directional heading. Thewavy camber line rib 62 of the present invention may still be describedas having a general arcing path across which it winds, and that thispath typically extends from an origination point near the leading edge28 and a trailing point near the trailing edge 29 of the airfoil 25. Itwill be appreciated that, in the case of a wavy camber line rib 62, itis this general arcing path that is roughly parallel to the camber lineof the airfoil 25.

Many known airfoil 25 configurations, such as the four-wall example ofFIG. 5 discussed above, include two camber line ribs 62. This type ofconfiguration may be described as having a pressure side camber line rib63 that resides nearer the pressure side outer wall 26, and a suctionside camber line rib 64 that resides nearer the suction side outer wall27. The present invention, as shown in FIG. 6, may includeconfigurations in which both the suction side camber line rib 64 and thepressure side camber line rib 63 are formed as wavy ribs. In alternativeembodiments, only one of these camber line ribs 62 may have a wavyprofile. It will be appreciated that the present invention may also beemployed in configurations having only a single camber line rib 62.

In airfoils 25 that include two camber line ribs 62, it will beappreciated that the pressure side camber line rib 63 and the suctionside camber line rib 64 define a center flow passage 40. The wavyprofile for each of the pressure side camber line rib 63 and the suctionside camber line rib 64 may be defined relative to the shape taken bysuccessive segments of the camber line rib 62 facing center flow passage40. That is, for example, relative to the central flow passage 40, thewavy profile of the camber line rib 62 may be described as including twosuccessive segments in which a first concave segment transitions to asecond convex segment. In an alternative embodiment, the wavy profilemay include four or more successive segments in which: a first concavesegment transitions to a second convex segment; the second convexsegment transitions to a third concave segment; and the third concavesegment transitions to a fourth convex segment.

According to an aspect of the present invention, the internal structureof an airfoil may include wavy ribs along the camber line direction ofthe airfoil. By making the camber line rib 62 into a spring in this way,the internal backbone of the airfoil may be made more compliant so thatperformance advantages may be achieved. In addition, the traverse ribsof the airfoil structure may be curved so to further soften the loadpath, as well as making more compliant connections with the ribs 62 andouter walls 26, 27 that they connect. Whereas standard linear ribdesigns experience high stress and low cyclic life due to the thermalfight between the internal cooling cavity walls and the much hotterouter walls, the present invention provides a spring-like constructionthat is better able to disburse stress concentrations, which, asprovided herein, may be used to improve the life of the component.

FIGS. 7 and 8 provide cross-sectional views of an airfoil having avortex inducing crossover passage 111 according to another aspect of thepresent invention. It will be appreciated that FIG. 7 shows a close-upof flow passages 40 on the suction side of an airfoil 25, while FIG. 8illustrates a similar view of flow passages 40 on the pressure side ofan airfoil with 5. FIG. 7 includes a more traditional configuration,while FIG. 8 illustrates an arrangement that includes a wavy camber linerib 62 having a wavy profile. As indicated, the crossover passage 111fluidly connects an inlet 120 formed in a flow passage 40 to an outlet117 formed in a second or downstream flow passage 40. It will beappreciated that such crossover passages 111 are typically are includedin internal configurations so to provide flow communication between flowpassages or are present due to casting necessities (i.e., the remnantsof support connectors used during the casting process, which typicallyare required to maintain a desired spatial relationship elongatedportions of the casting core). According to the present invention,crossover passages 111 of this type are formed having an angled orcanted configuration relative to the downstream flow passage, i.e., theflowpath such within which the outlet 117 is formed.

The canted configuration, according to the present invention, may bedescribed in several ways. For example, a center axis of the crossoverpassage 111 may be angled relative to a direction perpendicular to thesurface that surrounds the outlet 117 of the crossover passage 111. Inthis case, the canted configuration may include one in which the centeraxis of the crossover passage 111 and the direction perpendicular to thesurrounding surface of the outlet 117 define an acute angle that is atleast 20 degrees. Another way to describe this relationship, asillustrated in FIGS. 7 and 8, is that a linear crossover passage 111 isangled relative to the surface surrounding the outlet 117 of thecrossover passage 111. In this case, the crossover passage 111 and thesurface surrounding the outlet 117 of the crossover passage 111preferably form an angle 113 that is less than 70 degrees. In analternative embodiment, the crossover passage 111 and the surfacesurrounding the outlet 117 form an angle 113 that is less than 50degrees. More generally, the canted configuration also may be describedas one in which the crossover passage 111 is directed tangential to acenter axis of the second flow passage.

FIG. 9, which provides a longitudinal section of the airfoil 25,illustrates another aspect of the present invention. As indicated, thecrossover passage 111 includes a radial tilt, which, as used herein, isthe degree to which the crossover passage 111 is canted relative to apurely radial orientation. Given the orientation of the flow passages inFIG. 9, it will be appreciated that the radial direction is representedby the surfaces of the walls 26, 62. In preferred embodiments, theradial tilt of the canted configuration includes an angle 119 of atleast 20 degrees. In an alternative embodiment, the radial tilt of thecanted configuration includes an angle 119 of at least 40 degrees.

The crossover passages 111 may include a narrow configuration that issufficient to impinge a flow of pressurized coolant and aim the impingedflow toward a targeted region. It will be appreciated that impinging aflow of coolant in this way increases its cooling effectiveness. Thedownstream flow passage (i.e., the flow passage into which coolantpassing through the crossover passages 111 flows) may be described ashaving a number of radially extending sides. Some flow passages arerectangular in shape and include four such side, others may have more orless. The outlet 117 of the crossover passage 111 is formed on one ofthese sides, which may be referred to as the outlet side 116, whileanother of the sides is a targeted side, i.e., the side at which thecrossover passage 111 is aimed. In some cases, the side that includesthe outlet 117 and the targeted side are opposite sides or sides thatoppose each other across the downstream flow passage. In this case, thetargeted side is an opposite side 114. In a preferred embodiment, theside that includes the outlet 117 and the targeted inside are adjacentsides. In this case, the targeted side is an adjacent side 115. Asindicated, when the targeted side is an adjacent side 115, the crossoverpassage 111 may be positioned near that adjacent side.

In an alternative embodiment, as illustrated in both FIGS. 7 and 8, thetargeted side includes a turbulator 118. As one of ordinary skill in theart will appreciate, a turbulator 118 is an elongated protrusion that istypically rounded and includes steep sides. As indicated, the turbulator118 may be radially oriented and positioned on the targeted side. Itwill be appreciated that the turbulator 118 increases the surface areaof the inner wall and induces turbulent flow, which increases the heattransfer coefficient in this area. As illustrated, the turbulator 118also may be positioned relative to the canted crossover passage 111 soto enhance the formation of vortices, which is indicated by the flowarrows of both FIGS. 7 and 8.

As illustrated in FIG. 8, in a preferred embodiment, the crossoverpassage 111 may have a curved path. The curved path may inject thecoolant in such a way that it hugs closer to the walls of the flowpassage, which, for example, may be used advantageously to enhance theeffectiveness of turbulators 118, as indicated in FIG. 8. The curvatureof the path, generally, may be used to enhance the vortex induced in theflow passage 40.

The crossover passages 111 may be used to fluidly connect any of severaltypes of flow passages 40 that have been discussed the presentapplication. In one exemplary embodiment in which the targeted wall isan opposite side 114 (i.e., the side of the flow passage 40 that isopposite the outlet side 116), the outlet 117 is positioned on a camberline rib 62. In this case, the targeted wall is either a portion of adifferent camber line rib 62, or the targeted wall may be one of theouter walls, i.e., the pressure side outer wall 26 or the suction sideouter wall 27), which is the embodiments shown in FIGS. 7 and 8. Incases where the targeted side is an adjacent side 115 of the flowpassage 40, the outlet 117 preferably is positioned on a camber line rib62, while the targeted wall is a traverse rib 66. It will be appreciatedthat this may include pressure side traverse ribs 67, suction sidetraverse ribs 68, or center traverse ribs 69. Alternatively, the outlet117 may be positioned on a traverse rib 66, while the adjacent, targetedside of the flow passage 40 is portion of a camber line rib 62. In apreferred embodiment, either the outlet side 116 or the targeted wall isa camber line rib that has a wavy profile. The wavy profile may includeany of the configurations discussed above in relation to FIG. 6.

In operation, the crossover passage 111 of the present invention isconfigured in such a way to induce a vortex or non-radial flow pattern.This is accomplished by configuring the crossover passage 111 so that ithas a canted configuration, as defined above, relative to the flowpassage 40 into which it delivers coolant. It will be appreciated thatthe heat transfer coefficient of a cooling flow through a cavity can beenhanced by such vortices. For example, coolant having a smooth laminarflow through a flow passage typically becomes less effective as itconvects heat from the surrounding walls because the flow of coolantnear those walls becomes warmer and the temperature differential betweenit and the walls is reduced. The vortices and turbulent flow induced bythe crossover passages 111 and/or the turbulators 118 of the presentinvention disrupt such flow in a way that brings fresh, lowertemperature coolant from the middle of the flow passage to thesurrounding surfaces. Accordingly, aspects of the present vegan may beemployed to induce such vortices in areas requiring higher heat transfercoefficients. In this manner, flow passages may be tuned to reduceformation of thermal gradients in the airfoil. Because the effectivelife of a turbine blade is dependent upon the level of thermal gradientsand the forces/stresses that such temperature differences produce in thestructure, the configurations of crossover passages as described in thepreceding paragraphs may be used to thermally balance the airfoil andextend its useful life.

As one of ordinary skill in the art will appreciate, the many varyingfeatures and configurations described above in relation to the severalexemplary embodiments may be further selectively applied to form theother possible embodiments of the present invention. For the sake ofbrevity and taking into account the abilities of one of ordinary skillin the art, all of the possible iterations is not provided or discussedin detail, though all combinations and possible embodiments embraced bythe several claims below or otherwise are intended to be part of theinstant application. In addition, from the above description of severalexemplary embodiments of the invention, those skilled in the art willperceive improvements, changes and modifications. Such improvements,changes and modifications within the skill of the art are also intendedto be covered by the appended claims. Further, it should be apparentthat the foregoing relates only to the described embodiments of thepresent application and that numerous changes and modifications may bemade herein without departing from the spirit and scope of theapplication as defined by the following claims and the equivalentsthereof.

We claim:
 1. A turbine blade comprising an airfoil defined by a concaveshaped pressure side outer wall and a convex shaped suction side outerwall that connect along leading and trailing edges and, therebetween,form a radially extending chamber for receiving the flow of a coolant,the turbine blade further comprising: a rib configuration thatpartitions the chamber into radially extending flow passages thatinclude a first flow passage and a second flow passage; and a crossoverpassage that fluidly connects an inlet formed in the first flow passageto an outlet formed in the second flow passage; wherein the crossoverpassage comprises a canted configuration relative to the second flowpassage.
 2. The turbine blade according to claim 1, wherein the cantedconfiguration comprises one in which a center axis of the crossoverpassage is angled relative to a direction perpendicular to a surfacesurrounding the outlet of the crossover passage.
 3. The turbine bladeaccording to claim 2, wherein the canted configuration comprises one inwhich the center axis of the crossover passage and the directionperpendicular to the surface surrounding the outlet form an acute anglethat is greater than 20 degrees.
 4. The turbine blade according to claim1, wherein the crossover passage is approximately linear; wherein thecanted configuration comprises one in which the crossover passage isangled relative to a surface surrounding the outlet of the crossoverpassage.
 5. The turbine blade according to claim 4, wherein the cantedconfiguration comprises one in which the crossover passage and thesurface surrounding the outlet of the crossover passage form an anglethat is less than 70 degrees.
 6. The turbine blade according to claim 4,wherein the canted configuration comprises one in which the crossoverpassage and the surrounding surface of the outlet of the crossoverpassage form an angle that is less than 50 degrees.
 7. The turbine bladeaccording to claim 1, wherein the canted configuration comprises one inwhich the crossover passage is directed tangentially to a center axis ofthe second flow passage; wherein the crossover passage is curved.
 8. Theturbine blade according to claim 1, wherein the canted configuration ofthe crossover passage comprises a radial tilt.
 9. The turbine bladeaccording to claim 8, wherein the radial tilt of the cantedconfiguration comprises an angle of at least 20 degrees.
 10. The turbineblade according to claim 1, wherein the crossover passage comprises acurved profile.
 11. The turbine blade according to claim 1, wherein thesecond flow passage is defined by radially extending sides that include:a first side in which the outlet of the crossover passage is formed anda second side at which the crossover passage is aimed.
 12. The turbineblade according to claim 11, wherein the first side and the second sidecomprised opposite sides of the second flow passage.
 13. The turbineblade according to claim 11, wherein the first side and the second sidecomprised adjacent sides of the second flow passage.
 14. The turbineblade according to claim 13, wherein the second flow passage includes athird side that is opposite to the second side and adjacent to the firstside; wherein the outlet comprises a position on the first side that iscloser to the second side than the third side; and wherein the crossoverpassage comprises a curved profile in which the curvature is toward thefirst side.
 15. The turbine blade according to claim 11, wherein thesecond side comprises a turbulator, the turbulator comprising anapproximately radially oriented, elongated protrusion having steepsides.
 16. The turbine blade according to claim 12, wherein each of thefirst side and the second side comprises one of a pressure side outerwall, a suction side outer wall, and a camber line rib.
 17. The turbineblade according to claim 13, wherein the first side comprises a camberline rib and the second side comprises a traverse rib.
 18. The turbineblade according to claim 13, wherein the first side comprises a traverserib and the second side comprises a camber line rib.
 19. The turbineblade according to claim 13, wherein one of the first side and thesecond side comprises a segment of a camber line rib having a wavyprofile; and wherein the wavy profile includes one having at least oneback-and-forth “S” shape; wherein the turbine blade comprises a turbinerotor blade; and wherein the crossover passage comprises a narrowconfiguration that is configured to impinge a flow of pressurizedcoolant therethrough.
 20. A turbine blade comprising an airfoil definedby a concave shaped pressure side outer wall and a convex shaped suctionside outer wall that connect along leading and trailing edges and,therebetween, form a radially extending chamber for receiving the flowof a coolant, the turbine blade further comprising: a rib configurationthat partitions the chamber into radially extending flow passages thatinclude a first flow passage and a second flow passage; a crossoverpassage that fluidly connects an inlet formed in the first flow passageto an outlet formed in the second flow passage; wherein the crossoverpassage is linear and configured so to form an acute angle of at least20 degrees with a surface surrounding the outlet; wherein the secondflow passage is defined by radially extending sides that include: afirst side in which the outlet of the crossover passage is formed and asecond side at which the crossover passage is aimed; and wherein thesecond side includes a turbulator.